Methods and apparatus for supplying cooling fluid to turbine nozzles

ABSTRACT

A method enables a gas turbine engine to be operated. The method comprises supplying cooling fluid into a manifold ring that includes a plurality of distribution ports defined by a sidewall connected by a radially inner wall, channeling the cooling fluid circumferentially through the manifold ring and through at least one distribution port that is defined by a wall that extends arcuately across at least one turbine nozzle, and discharging cooling fluid from the distribution ports radially inwardly towards the at least one turbine nozzle positioned radially inward from the manifold ring.

BACKGROUND OF THE INVENTION

[0001] This invention relates generally to gas turbine engine nozzlesand more particularly, to methods and apparatus for supplying coolingfluid to turbine nozzles.

[0002] Gas turbine engines include combustors which ignite fuel-airmixtures which are then channeled through a turbine nozzle assemblytowards a turbine. At least some known turbine nozzle assemblies includea plurality of nozzles arranged circumferentially. At least some knownturbine nozzles include a plurality of circumferentially-spaced hollowairfoil vanes coupled by integrally-formed inner and outer bandplatforms. More specifically, the inner band forms a portion of theradially inner flowpath boundary and the outer band forms a portion ofthe radially outer flowpath boundary

[0003] At least some known turbine nozzle airfoil vanes are hollow andinclude a cavity defined therein. At least some known airfoil vanecavities are partitioned into a leading edge cooling passage, a centerpassage, and a trailing edge passage. Cooling air is supplied to amanifold that extends circumferentially within the engine and around theturbine nozzle outer bands. The airflow is directed radially inwardlythrough a plurality of distribution ports that are formed integrallywith the manifold. Specifically, the distribution ports arecircumferentially-spaced about the manifold such to facilitate supplyingcooling air to a respective turbine nozzle vane. More specifically,known distribution ports are defined by a substantially planar radiallyinner surface that extends across the port to an annular sidewall. Tofacilitate distributing the cooling airflow across the turbine nozzleairfoil vane cooling passages, known distribution ports include aplurality of openings that extend through the distribution portsidewall. Accordingly, the openings are oriented approximately ninetydegrees from the distribution port radially inner surface.

[0004] During operation, airflow channeled to the distribution ports isforced radially inward towards the nozzle cavities through thedistribution port openings. The airflow entering the distribution portsimpinges against the radially inner surface of the ports and is changedin flow direction through the port openings. However, as the airflowchanges flow direction, turbulence and pressure losses are generated.Moreover, the turbulence and pressure losses may adversely effectcooling airflow supplied to the airfoil vane cavity center coolingpassage. Over time continued operation with decreased cooling of thecenter cooling passage, may limit a useful life of the turbine nozzle.

BRIEF SUMMARY OF THE INVENTION

[0005] In one aspect, a method for operating a gas turbine engine isprovided. The method comprises supplying cooling fluid into a manifoldring that includes a plurality of distribution ports defined by asidewall connected by a radially inner wall, channeling the coolingfluid circumferentially through the manifold ring past into at least onedistribution port that is defined by a radially inner wall that extendsarcuately across the distribution port, and discharging cooling fluidfrom the manifold ring radially inwardly towards a plurality of turbinenozzles positioned radially inward from the manifold ring.

[0006] In another aspect, a turbine nozzle assembly for a gas turbineengine is provided. The nozzle assembly includes a turbine nozzle, and amanifold ring. The turbine nozzle includes an outer band, an inner band,and plurality of airfoil vanes coupled together by the outer and innerbands. Each airfoil vane is hollow and defines a cavity therein. Themanifold ring extends circumferentially around the turbine nozzle forchanneling cooling fluid radially inwardly into each airfoil vanecavity. The manifold ring includes a radially outer wall and a radiallyinner wall coupled together by a pair of sidewalls. At least a portionof the radially inner wall extends arcuately between the pair ofsidewalls.

[0007] In a further aspect, a gas turbine engine including a turbinenozzle assembly is provided. The turbine nozzle assembly includes atleast one turbine nozzle, and a manifold ring. The at least one turbinenozzle includes an outer band, an inner band, and plurality of airfoilvanes coupled together by the outer and inner bands. Each of the airfoilvanes is hollow and defines a cavity therein. The manifold ring extendscircumferentially within the gas turbine engine and is radially outwardfrom the at least one turbine nozzle for channeling cooling fluidradially inward into each airfoil vane cavity. The manifold ringincludes at least one distribution port defined by a radially inner wallthat extends arcuately across the distribution port between a sidewall.

BRIEF DESCRIPTION OF THE DRAWINGS

[0008]FIG. 1 is a schematic illustration of a gas turbine engine;

[0009]FIG. 2 is a perspective view of an exemplary turbine nozzle thatmay be used with the gas turbine engine shown in FIG. 1;

[0010]FIG. 3 is a cross-sectional view through a plurality of nozzlevanes that may be used with the turbine nozzle shown in FIG. 2;

[0011]FIG. 4 is an enlarged side cross-sectional view of a distributionport that may be used with a manifold ring shown in FIG. 2 to supplycooling fluid to the turbine nozzle shown in FIG. 2; and

[0012]FIG. 5 is an enlarged side cross-sectional view of an alternativeembodiment of a distribution port that may be used to supply coolingfluid to a turbine nozzle, such as the turbine nozzle shown in FIG. 2.

DETAILED DESCRIPTION OF THE INVENTION

[0013]FIG. 1 is a schematic illustration of a gas turbine engine 10including a low pressure compressor 12, a high pressure compressor 14,and a combustor 16. Engine 10 also includes a high pressure turbine 18and a low pressure turbine 20. Compressor 12 and turbine 20 are coupledby a first shaft 22, and compressor 14 and turbine 18 are coupled by asecond shaft 21.

[0014] In operation, air flows through low pressure compressor 12 andcompressed air is supplied from low pressure compressor 12 to highpressure compressor 14. The highly compressed air is delivered tocombustor 16. Airflow from combustor 16 exits combustor 16 and drivesturbines 18 and 20, and then exits gas turbine engine 10.

[0015]FIG. 2 is a perspective view of an exemplary known turbine nozzle50 including a manifold ring 51 that may be used with a gas turbineengine, such as gas turbine engine 10. FIG. 3 is a cross-sectional viewthrough nozzle 50. FIG. 4 is an enlarged side cross-sectional view of adistribution port 140 that may be used to supply cooling fluid toturbine nozzle 50. In the exemplary embodiment, nozzle 50 includes aplurality of circumferentially-spaced airfoil vanes 52 coupled togetherby an arcuate radially outer band or platform 54, and an arcuateradially inner band or platform 56. More specifically, in the exemplaryembodiment, each band 54 and 56 is integrally-formed with airfoil vanes52, and nozzle 50 includes at least one airfoil vane 52. Alternatively,nozzle 50 includes a plurality of vanes 52.

[0016] In the exemplary embodiment, airfoil vanes 52 are substantiallyidentical and each nozzle includes a leading airfoil vane 76 and atrailing airfoil vane 78. Each individual vane 52 includes a firstsidewall 80 and a second sidewall 82. First sidewall 80 is convex anddefines a suction side of each airfoil vane 52, and second sidewall 82is concave and defines a pressure side of each airfoil vane 52.Sidewalls 80 and 82 are joined at a leading edge 84 and at anaxially-spaced trailing edge 86 of each airfoil vane 52. Morespecifically, each airfoil trailing edge 86 is spaced chordwise anddownstream from each respective airfoil leading edge 84.

[0017] First and second sidewalls 80 and 82, respectively, extendlongitudinally, or radially outwardly, in span from radially inner band56 to radially outer band 54. Additionally, first and second sidewalls80 and 82, respectively, define a cooling cavity 90 within each airfoilvane 52. More specifically, cavity 90 is bounded by an inner surface 92and 94 of each respective airfoil sidewall, and extends through eachband 54 and 56.

[0018] Each cooling cavity 90 is coupled to a plurality of coolingpassages 100 defined therein. More specifically, in the exemplaryembodiment, passages 100 include a leading edge cooling passage 120, atrailing edge cooling passage 122, and a center cooling passage 124 thatextends therebetween.

[0019] Manifold ring 51 is annular and extends circumferentially withinengine 10, such that manifold ring 51 is radially outward from turbinenozzles 50. Specifically, manifold ring 51 connects to casing 74, andchannels cooling air to each distribution port 140. Each port 140 isgenerally centered with respect to, and radially outward from, eachturbine nozzle outer band 56. Manifold ring 51 is hollow and includes aradially outer end 138 and a radially inner wall 134 coupled together bya pair of radial sidewalls 136.

[0020] Manifold ring 51 includes a plurality of circumferentially spaceddistribution ports 140 that extend radially inward. As is known in theart, each distribution port 140 is coupled to a cooling source thatprovides cooling fluid for cooling of nozzles 50. In one embodiment,each port 140 receives cooling air from a compressor, such as compressor14 (shown in FIG. 1).

[0021] Within each distribution port 140 is a tube 132, known as aspoolie, that is substantially centered and extends radially inward, toconnect with a manifold port 300. More specifically, in the exemplaryembodiment, each spoolie 132 is positioned at least partially radiallybetween each distribution port 140 and manifold port 300, and extendsbetween each port 140 and nozzle manifold port 300. Each manifold port300 is also connected to a cover plate 201 which is coupled to radialouter faces of nozzle 50. More specifically, cover plate 201 extendsacross nozzle 50 and includes a contoured bottom surface 152 thatfacilitates directing cooling fluid into nozzle 50. In the exemplaryembodiment, cover plate bottom surface 152 is substantiallysemi-spherical.

[0022] A plurality of cooling openings 154 extend through bottom 152.More specifically, in the exemplary embodiment, openings 154 areidentically sized and are spaced axially and circumferentially acrosscover plate 201. Openings 154 control the flow of cooling fluid frommanifold port 300. More specifically, as described in more detail below,the size and relative location of openings 154 is variably selected tofacilitate providing a substantially even flow distribution of coolingfluid to airfoil vane cavity cooling passages 100.

[0023] During operation, as hot combustion gases flow through nozzles50, an operating temperature of nozzles 50 is increased. Cooling fluidsupplied to cooling manifold ring 51 is circulated in a circumferentialdirection before being channeled radially inwardly into distributionport 140. The cooling fluid is then forced radially inwardly throughdistribution port 140 towards spoolie 132, wherein the fluid ischanneled into manifold port 300. The cooling fluid is then dischargedtowards cover plate 201 wherein because cooling plate bottom 152contoured radially inward towards nozzles 50, rather than beingsubstantially planar, airflow impinging on port bottom 152 is notreflected in a reverse effect such that airflow buffeting is generatedwithin port 140. Rather, the contour of bottom 152 facilitates improvingthe air flow distribution supplied to nozzle airfoil cavity 90.

[0024]FIG. 5 is an enlarged side cross-sectional view of an alternativeembodiment of a distribution port 200 that may be used to supply coolingfluid to a turbine nozzle, such as the turbine nozzle 50 (shown in FIG.2). Cooling distribution port 200 is substantially similar todistribution port 140 shown in FIG. 4, and components in distributionport 200 that are identical to components of cooling distribution port140 are identified in FIG. 5 using the same reference numerals used inFIG. 4. Accordingly, distribution port 200 is substantially circular andextends radially from manifold ring 51.

[0025] Distribution port 200 also includes a plurality ofcircumferentially-spaced distribution ports 210 that extend radiallyinward to connect with spoolie 132 which is coupled to manifold port300. Manifold port 300 is coupled to a cover plate 214 which extendsacross nozzle 50 and includes a contoured inlet, which in the exemplaryembodiment, is a dimpled contoured bottom 212. Because bottom 212 iscontoured, rather than planar, bottom 212 facilitates directing coolingfluid into nozzle 50. More specifically, in the exemplary embodiment,each distribution port 200 is substantially centered with respect to,and radially outward from, each turbine nozzle 50, such that each port200 is at least partially positioned axially within each segment. As isknown in the art, distribution port 200 is coupled to a cooling sourcethat provides cooling fluid to nozzles 50. In one embodiment, port 200receives cooling air from a compressor, such as compressor 14 (shown inFIG. 1).

[0026] Each distribution port 200 is substantially circular and iscoupled to spoolie 132 and manifold port 300. In the exemplaryembodiment, bottom 212 is formed integrally with cover plate 214, andextends arcuately across nozzle 50. More specifically, bottom 212 iscontoured or bows radially outwardly away from nozzles 50 and towardsdistribution port 200. Additionally, in the exemplary embodiment,sidewalls 216 extend substantially perpendicularly from cover plate 214.In an alternative embodiment, sidewalls 216 are not orientedsubstantially perpendicular with respect to cover plate 214.

[0027] A plurality of cooling openings 154 extend through sidewall 216.More specifically, in the exemplary embodiment, openings 154 areidentically sized and are spaced axially and circumferentially acrossdistribution port 200. Openings 154 control the flow of cooling fluidfrom distribution port 200 into each airfoil vane cavity 90. Morespecifically, the size and relative location of openings 154 is variablyselected to facilitate providing a substantially even flow distributionof cooling fluid to airfoil vane cavity cooling passages 100

[0028] The above-described turbine nozzle includes a plurality ofcircumferentially-spaced distribution ports that extend radiallyinwardly from the manifold ring towards each turbine nozzle. Morespecifically, each distribution port includes a radially inner surfaceor bottom portion that extends across the port between the sidewalls.Because the bottom portion is arcuately contoured between the sidewalls,cooling fluid impinging on the bottom portion is not reflected radiallyoutwardly to generate buffeting within the distribution. Rather, thearcuate surface facilitates reducing the velocity of the cooling fluidsuch that pressure losses associated with the cooling fluid turningthrough the distribution port openings are facilitated to be reduced incomparison to known distribution ports that include substantially planarradially inner surfaces. As a result, cooling fluid is supplied to theturbine nozzle in a flow distribution pattern that enables each coolingpassage defined within the turbine nozzle airfoil vanes to receiveapproximately the same volume of cooling fluid. Accordingly, the turbinenozzle airfoil vanes are operable at a reduced operating temperature,which facilitates extending the durability and useful life of theturbine nozzles.

[0029] Exemplary embodiments of turbine nozzles are described above indetail. The nozzles are not limited to the specific embodimentsdescribed herein, but rather, components of each turbine nozzle may beutilized independently and separately from other components describedherein.

[0030] While the invention has been described in terms of variousspecific embodiments, those skilled in the art will recognize that theinvention can be practiced with modification within the spirit and scopeof the claims.

What is claimed is:
 1. A method for operating a gas turbine engine, saidmethod comprising: supplying cooling fluid into a manifold ring thatincludes a plurality of distribution ports defined by a sidewallconnected by a radially inner wall; channeling the cooling fluidcircumferentially through the manifold ring and through at least onedistribution port that is defined by a wall that extends arcuatelyacross at least one turbine nozzle; and discharging cooling fluid fromthe distribution ports radially inwardly towards the at least oneturbine nozzle positioned radially inward from the manifold ring.
 2. Amethod in accordance with claim 1 wherein discharging cooling fluid fromthe distribution ports radially inwardly further comprises dischargingcooling fluid from at least one distribution port through a plurality ofopenings formed within the at least one distribution port and definedwithin the wall that extends arcuately across the at least one turbinenozzle.
 3. A method in accordance with claim 1 wherein channeling thecooling fluid circumferentially through the manifold ring and through atleast one distribution port further comprises channeling the coolingfluid through the at least one distribution port and past a radiallyinner wall that is contoured radially inward towards the at least oneturbine nozzle.
 4. A method in accordance with claim 1 whereinchanneling the cooling fluid circumferentially through the manifold ringand through at least one distribution port further comprises channelingthe cooling fluid through the at least one distribution port and past aradially inner wall that is contoured radially outward from the at leastone turbine nozzle.
 5. A method in accordance with claim 1 whereinchanneling the cooling fluid circumferentially through the manifold ringand through at least one distribution port further comprises channelingthe cooling fluid through a plurality of circumferentially-spacedsubstantially circular distribution ports that extend radially inwardtowards a plurality of turbine nozzles.
 6. A turbine nozzle assembly fora gas turbine engine, said nozzle assembly comprising: a turbine nozzlecomprising an outer band, an inner band, and plurality of airfoil vanescoupled together by said outer and inner bands, each said airfoil vaneis hollow and defines a cavity therein; and a manifold ring extendingcircumferentially around said turbine nozzle for channeling coolingfluid radially inwardly into each said airfoil vane cavity, saidmanifold ring comprising a radially outer wall and a radially inner wallcoupled together by a pair of sidewalls, at least a portion of saidradially inner wall extends arcuately between said pair of sidewalls. 7.A turbine nozzle assembly in accordance with claim 6 wherein at least aportion of said radially inner wall is contoured radially inward towardseach said turbine nozzle cavity.
 8. A turbine nozzle assembly inaccordance with claim 6 wherein at least a portion of said distributionport radially inner wall is contoured radially outward away from eachsaid turbine nozzle cavity.
 9. A turbine nozzle assembly in accordancewith claim 6 wherein said manifold ring is hollow and defines a cavitytherein, said manifold ring further comprises a plurality of openingsextending through at least one of said pair of sidewalls and saidradially inner wall, said plurality of openings in flow communicationwith said manifold ring cavity for discharging cooling fluid therefrom.10. A turbine nozzle assembly in accordance with claim 9 wherein saidturbine nozzle airfoil vane cavity contains a plurality of coolingpassages defined therein, said plurality of manifold cooling openings inflow communication with said plurality of cooling passages.
 11. Aturbine nozzle assembly in accordance with claim 6 wherein said manifoldring facilitates enhancing cooling fluid flow distribution into saidturbine nozzle airfoil vane cavity.
 12. A turbine nozzle assembly inaccordance with claim 6 wherein said manifold ring radially inner wallfurther comprises a plurality of circumferentially-spaced substantiallyspherical projections extending radially inward towards each saidturbine nozzle airfoil vane cavity.
 13. A gas turbine engine comprisinga turbine nozzle assembly comprising at least one turbine nozzle and amanifold ring, said at least one turbine nozzle comprising an outerband, an inner band, and plurality of airfoil vanes coupled together bysaid outer and inner bands, each of said airfoil vanes is hollow anddefines a cavity therein, said manifold ring extends circumferentiallywithin said gas turbine engine and is radially outward from said atleast one turbine nozzle for channeling cooling fluid radially inwardinto each said airfoil vane cavity, said manifold ring comprises atleast one distribution port defined by a radially inner wall thatextends arcuately across said at least one distribution port.
 14. A gasturbine engine in accordance with claim 13 wherein said manifold ring ishollow and defines a cavity therein, said manifold ring furthercomprises a plurality of openings extending through at least a portionof said distribution port radially inner wall, said plurality ofopenings in flow communication with said manifold ring cavity.
 15. A gasturbine engine in accordance with claim 14 wherein each said vane cavitycontains a plurality of cooling passages defined therein, said pluralityof manifold ring cooling openings in flow communication with said vanecavity plurality of cooling passages.
 16. A gas turbine engine inaccordance with claim 14 wherein at least a portion of said distributionport radially inner wall is contoured radially inwardly towards eachsaid turbine nozzle airfoil vane cavity.
 17. A gas turbine engine inaccordance with claim 14 wherein at least a portion of said distributionport radially inner wall is contoured radially outwardly away from saidturbine nozzle airfoil vane cavity.
 18. A gas turbine engine inaccordance with claim 14 wherein said manifold ring facilitatesenhancing cooling fluid flow distribution into each said turbine nozzleairfoil vane cavity.
 19. A gas turbine engine in accordance with claim14 wherein said manifold ring radially inner wall further comprises aplurality of substantially circular distribution ports extendingradially inward towards each said turbine nozzle airfoil vane cavities.